This invention relates to gas turbine engines, and more particularly, to gas turbine engines employing off-axis pulse detonation engines and a reverse flow turbine.
Gas turbine engines, such as those used in aircraft engine applications, typically have three main stages. A compressor stage, a combustion stage and a turbine stage. The compressor stage feeds compressed air flow into the combustion stage, where the compressed flow is combusted with fuel to raise the temperature of the mixture. Subsequent thereto, the combusted mixture enters the turbine stage, where the mixture expands, thus driving the turbine. As the combusted mixture exits the turbine stage, and ultimately the engine, thrust is produced.
However, standard combustion systems use energy release in the combustion process to raise only the temperature of the compressed flow. This represents a very inefficient process of the thermodynamic cycle. No useful work is accomplished, as there is an overall pressure drop in the combustion stage.
In recent years pulse detonation combustion systems have been developed. These systems use compressed air mixed with a fuel to create rapid combustion and compression of a fuel/air mixture. The rapid combustion and compression produce detonation waves, which not only increase the gas temperature, but also provide a meaningful increase in pressure.
Thus, when pulse detonation systems are employed in a gas turbine engine, the increase in pressure can be used to reduce the number of stages or airfoils in the compression stage, and/or can be used to raise the overall cycle pressure ratio. However, one difficulty with the incorporation of pulse detonation systems into gas turbine engines is that the required length of the overall configuration can be undesirable. Current pulse detonation systems are typically longer than existing typical combustor stages. As such, the added length, placed between typical compressor and turbine stages, can be unattractive or impractical for most gas turbine engine applications.